1. Field of the Invention
This invention relates generally to a load carrying tubular member and, more particularly to a tubular member that has been wound with specific cross-section fibers in a controlled orientation to optimally carry the load applied to the member, where a number of load carrying members can be assembled to cooperate in forming a body of revolution.
2. Description of the Prior Art
Over the past two decades, the use of fiber composite materials in aircraft structures has gained popularity. As a result, modem air frames incorporate structural components made of composite materials to form aircraft wing structures, rotor blades, fuselage segments and the like as substantial weight savings can be achieved due to the superior strength-to-weight ratio of fiber composite materials as compared with the conventional materials of aircraft construction such as metal alloys. By replacing structural components previously formed of metal alloys with similar versions of the same component formed of composite material, a respective weight savings in the order of 25 to 30 percent is generally considered to be achievable.
In general, composites include a reinforcing material suspended in a xe2x80x9cmatrixxe2x80x9d material that stabilizes the reinforcing material and bonds it to adjacent reinforcing materials.
Composite parts are usually molded, and may be cured at room conditions or at elevated temperature and pressurized for greater strength and quality.
Most of the composites used in aircraft structures comprise of filament reinforcing material embedded in a polymer matrix. A primary advantage associated with the use of filament composites is that their structural properties may be tailored to the expected loads in different directions. Contrary to metals which have the same material properties in all directions, filament composites are strongest in the direction the fibers are running. If a structural element such as a spar is to carry substantial load in only one direction, all the fibers can be oriented in that direction. This characteristic of filament composite provides for exceptional strength-to-weigh ratios and offers a tremendous weight savings opportunity to structural designers.
When fibers are aligned in only one direction, the resulting structure has maximum strength in that direction, and has little strength in other directions. Therefore, multiple layers or xe2x80x9cpliesxe2x80x9d having fibers aligned in different directions with respect to one another are combined in a desired arrangement to provide combined strength along the principal axis as well as off-axis directions. As such, fibers oriented at 45xc2x0 degree angles with the principle axis provide strength in two directions. For this reason, the 45xc2x0 orientation is frequently used in structure that must resist torque. By utilizing permutations of this design philosophy to provide alternate plies of fibers at 0xc2x0, 45xc2x0, and 90xc2x0 orientations the structural designer can obtain virtually any combination of tensile, compression, and shear strength in desired directions.
Common forms of fiber used in the production of composite structures include unidirectional tape, unidirectional fabric and bidirectional fabric. Unidirectional tape typically comes pre-impregnated with matrix material and is customarily provided on large rolls which can then be placed in a mold by hand or by robotic tape-laying machines. Similarly, bidirectional fabrics, having fibers running at 0 and 90 degrees, or unidirectional fabrics having fibers running in one direction may also be provided on large rolls pre-impregnated with matrix material. In another form of composite, individual filaments are wound around plugs or mandrels to form desired structural shapes. By way of background, the mandrels duplicate the inner skin of the structure or the inner surface of the structure. This technique is known as filament wound construction.
In addition to the form of fiber used in the production of composite structures, there are a number of fiber and matrix combinations which can be employed to provide desired structural properties of the resulting aircraft components. Fiberglass embedded in an epoxy-resin matrix has been used for years for nonstructural components such as radomes and minor fairings. It is worthy of noting, however, that while fiberglass-epoxy has relatively good strength characteristics, its relatively low strength to weight ratio prevents its use in highly loaded structure. Additional material combinations which have eliminated this condition include: boron fibers used in combination with an epoxy matrix; aramid fibers (known as Kevlar) used in combination with an epoxy matrix, and graphite fibers used in combination with an epoxy matrix.
The United States military has been quick to incorporate fiber composite based structural components in its high-performance military aircraft. For example the F-16 utilizes graphite-epoxy composite material to form the horizontal and vertical tail skins. Similarly, graphite-epoxy composite material is utilized in the FA-18 where such material forms the wing skins, the horizontal and vertical tail skins, the fuselage dorsal cover, the avionics bay door, the speed brake, and many of the control surfaces. The AV-8B employs composite materials even more extensively. In the AV-8B almost the entire wing, including the skin and substructure, is made of graphite-epoxy composite material with such material comprising approximately 26% of the total aircraft structural weight.
While composite materials have played an important role in reducing the overall structural weight of modem air frames, it should be noted that the basic design and layout of primary load carrying components contained within these structures has remained relatively the same. For example, a conventional aircraft wing structure consists of individual components such as spars, ribs, stringers and skin sections joined in combination to provide an integrated load carrying body which is capable of reacting to aerodynamic forces encountered during flight. As a result, individual spars, ribs, stringers and skin sections are specifically sized and oriented relative to one another so as to provide an optimized structural assembly designed to efficiently carry localized stresses generated by the combined effects of lift, drag, wind gusts, and acceleration loads which interact with surface of the wing or other airframe components.
In order to take advantage of weight savings opportunities afforded by the use of lighter weight materials, individual spars, ribs, stringers and skin sections previously formed from metal alloys have been replaced by similar components formed of fiber composite material. Frequently, these lighter weight components incorporate a xe2x80x9csandwichxe2x80x9d style construction having two face sheets, or skins, made of fiber composite material which are bonded to and separated by a core. Typically, sandwich structures are formed with fiberglass-epoxy or graphite-epoxy skins which are bonded with adhesive to a phenolic honeycomb or rigid foam core wherein the skins carry tension and compression loads due to bending and the core carries shear loads as well as the compression loads perpendicular to the skins.
Unfortunately, manufacturing complexity and related labor cost associated with the assembly of numerous individual components, joined together to form an integrated load carrying structure, still remains. For example, conventional airframe construction techniques employ the use of elaborate jig fixtures designed to hold individual component parts in relative alignment during assembly to ensure proper component installation. In addition, drill templates are utilized to locate and drill fastener holes through mating pieces of structure to accommodate bolts or rivets used to mechanically join components together. These construction techniques are time consuming and require a great deal of dimensional precision because an improper installation of structural components may create a weakened resulting structure. Furthermore, the utilization of mechanical fasteners significantly contributes to overall structural weight. It is therefore generally desirable to minimize the number of mechanical joints in a structure in order to minimize both its weight and manufacturing cost while ensuring structural integrity. Integrally formed fiber composites structures have an important advantage over complicated structural assemblies in this respect, since large one-piece components are readily produced.
What has been needed and heretofore unavailable is a one-piece structure which is integrally formed as a unitary body and which is optimized to efficiently carry localized stresses developed from the complex interaction of static and aerodynamic forces encountered during all aspects of aircraft operation. The present invention satisfies these needs.
The present invention is directed to integrally stiffened load carrying structures comprising of a plurality of elongated thin-walled triangular tubes placed co-extensively in a complementary side-by-side fashion to form at least a portion of the wall of a hollow core having a desired external contour. Integral skins forming the external and internal surfaces of the core cooperate therewith to provide an integrally formed, unitary load carrying body of xe2x80x9csandwichxe2x80x9d style construction.
Upon the application of external forces to the structure, adjacent triangular tubes forming the core cooperate to react loads about the perimeter of the structure. Similarly, adjacent tubes forming an internal support member cooperate to transfer loads from one side of the structure to the other. It will be appreciated that the present invention is capable of providing various load carrying cross-sections. Therefore, the cross-sectional geometry of the load carrying body can be specifically designed to provide a desired external contour which is capable of reacting expected external forces applied thereto.
This structure can be formed by, but is not limited to, extrusion, casting, diffusion bonding, the controlled deposition of material at the atomic level, and filament winding. With regard to the controlled deposition, a controlled deposition method such as Laser-assisted Chemical Vapor Deposition (LCVD) process may be used. Of course, other methods known to one of ordinarily skilled in the art may also be used.
By utilizing well-known filament winding techniques, the material properties of each tube can be specifically tailored to react localized stresses generated from the application of external forces upon the structure. In general, a triangular tube is formed with multiple layers or xe2x80x9cpliesxe2x80x9d of composite material having fibers aligned in different directions. The plies of composite material are arranged with respect to one another to provide a structural element which is capable of reacting to forces in multiple directions. By utilizing alternate plies of fibers oriented at between 0xc2x0 and 90xc2x0 orientations relative to the longitudinal axis of the structure, each individual tube will be capable of reacting tensile, compression and shear stress from multiple directions. It will be appreciated that by tailoring the load carry capability of the individual tubes to suit the loads they are expected to encounter, a lightweight, efficient, load carrying structure may be produced.
It is also envisioned that the skins surrounding the internal and external surfaces of the shell and internal support member may be formed with filament wound fiber composite material. Like the construction of the individual triangular tubes discussed above, filament winding techniques may be utilized to tailor the load carrying properties of the skin. By providing layers of composite material having fibers running parallel to to the longitudinal axis of the structure, skins suited for carrying localized stresses resulting from the application of longitudinal bending loads may be produced. Likewise, by incorporating layers of composite material having fibers oriented at between 0xc2x0 and 90xc2x0 to the longitudinal direction, the skins may also have the ability to react shear stresses resulting from torsional loading of the structure.
In order to design and fabricate integrally stiffened load carrying composite structures embodying the present invention, an estimation of the external forces which will be reacted by the proposed structure must be determined. This estimation requires a thorough understanding of the loading environment and operating conditions that the proposed design is expected to experience. Based upon these expected loading characteristics, the geometry of the proposed design can be used to resolve these forces and moments into resulting localized stresses. Individual structural components can then be appropriately sized and designed to efficiently carry these expected stresses.
Once the localized stresses are known, individual components which form the load carrying structure can be fabricated. The process of building up individual fiber reinforced skins and tubular elements is essentially a three-dimensional strengthening process. By utilizing filament winding techniques, fibers pre-impregnated with matrix material are wound under controlled tension to thereby precisely arrange multiple layers of fiber on a shaped mandrel surface.
From a structural design perspective, the tubular elements cooperating to form the load carrying shell are necessarily required to react stresses generated from more than one direction as resultant forces are applied to the structure from different directions. For example, a wing structure must be designed in such a way to efficiently react lifting forces and associated bending moments, frontal loads associated with aerodynamic drag and impulsive forces associated with wind gusts. Therefore, an important aspect of forming each individual tubular element is to orient the fibers along the mandrels in appropriate directions and proportions to form a composite structure having the desired mechanical properties suitable to carry anticipated localized stresses. While the winding process must produce the desired shape of each tubular element, in the ideal case, fibers will be aligned with the trajectories of principal stresses and will be concentrated in direct proportion to the local magnitude of stress.
After the individual triangular mandrels have been wound with an appropriate combination of fiber, they are placed together side-by-side in a geometrically complimentary fashion about appropriately shaped pre-wound mandrels to form the load carrying structure having a predetermined external contour. Additional fiber is then wound about the exterior of the assembly to provide a skin surrounding the exterior surface of the structure. The assembly is then placed into a mold having mold faces shaped to desired external contour of the structure. For most applications, this process eliminates the need for vacuum bagging and autoclaves. Temperature and pressure are employed by the mold to cure the composite, thereby bonding the skins and triangular tubes together. After the structure has cured, the individual mandrels are removed from the structure to provide an integrally formed, unitary load carrying body.
It will be appreciated that, by way of example and not of limitation, the present invention is capable of providing integrally stiffened aircraft wing structures, rotor blades, fuselage segments and the like, having a reinforced load carrying shell formed integral to an underlying support member, such as an X-shaped spar or strut. The skin, reinforced, shell and underlying internal support member thereby cooperate to carry static and aerodynamic forces encountered all aspects of aircraft operation. As a result of this novel method of construction, the need for individual stringers, ribs, spars, and skin sections typically used in combination to form conventional aircraft structures is eliminated.